The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
Advanced engine cycles require the combustor section to operate at relatively high compressor exit temperatures. A survey of typical flight envelopes often reveals that high compressor exit temperatures exist with reduced supply pressure at high altitude. These operational conditions may result in relatively high convection and radiation heat loads.
The combustor section typically includes a double wall assembly with an outer shell that is lined with heat shields, which are often referred to as floatwall liner panels, attached to the outer shell with studs and nuts. In certain arrangements, dilution holes through the double wall combustor direct cooling air for dilution of the combustion gases. In addition to the dilution holes, the outer shell typically includes numerous relatively smaller air impingement holes to direct cooling air between the floatwall panels and the outer shell of the double wall combustor to impingement cool the liner panels. This cooling air then exits effusion holes through the liner panels to form a cooling air film on a hot side of the liner panels that serves as a barrier to facilitate reduction of thermal damage.
The combustor liner panels may be subject to distress that varies both axially and circumferentially due to the complex turbulent currents of combustion products and dilution air.